Impingement cooling of combustor liners

ABSTRACT

A gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine. The double-walled liner may extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. The plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end. The at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.

TECHNICAL FIELD

The present disclosure relates generally to systems and methods of impingement cooling a combustor liner of a gas turbine engine.

BACKGROUND

Combustor liners of gas turbine engines are exposed to high temperatures of combustion and therefore require cooling. A type of combustor liner, called a double-walled liner, includes an inner liner that encloses a volume where combustion occurs and an outer liner that surrounds the inner liner. An annular space between the inner liner and the outer liner assists in the cooling of the liner. There are various methods that are employed to cool combustion liners during operation of the engine. These method include film cooling and jet impingement cooling. In film cooling, air in the annular space is directed into the combustor through holes in the inner liner to mix with the hot combustion gases within. The air absorbs the heat from the inner liner as it flows therethrough. In jet impingement cooling, air jets impinge upon and cool the back surface of the inner liner. These air jets may be directed to the back surface of the inner liner through an array of holes on the outer liner. After impinging on the back surface of the inner liner, the spent cooling air flows downstream through the annular space. This spent air flow, called cross-flow, is known to degrade the cooling ability of downstream air jets.

U.S. Patent Application No. 2008/0271458 to Ekkad et al. (the '458 publication) describes an impingement cooled liner with ports extending from the outer liner to the inner liner to reduce the effects of cross-flow. While the extended ports of the '458 publication may reduce the effects of cross-flow, they may have limitations. For instance, dimensional changes during operation of the turbine engine may force portions of the inner liner against the extended ports preventing air flow therethrough. The systems and methods of the current disclosure are directed to overcoming one or more of the problems set forth above.

SUMMARY

In one aspect, a gas turbine engine is disclosed. The gas turbine engine may include an impingement cooled double-walled liner, having an inner liner and an outer liner, disposed around a combustion space of the turbine engine. The double-walled liner may extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. The plurality of nozzles may be arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end. The at least one nozzle of the plurality of nozzles may include multiple air holes at the second end.

In another aspect, a method of impingement cooling a double-walled combustor liner of a gas turbine engine is disclosed. The double-walled liner may extend from an upstream end to a downstream end and include an inner liner and an outer liner positioned radially outwards the inner liner. The method may include combusting a fuel in a combustor of the gas turbine engine, and directing cooling air through a plurality of nozzles that extend radially inwards through the outer liner to impinge upon and cool the inner liner. The cooling air may be directed such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end. The cooling air directed through at least one nozzle of the plurality of nozzles may exit the at least one nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.

In yet another aspect, a gas turbine engine is disclosed. The gas turbine engine may include an impingement cooled double-walled liner. The double-walled liner may include an inner liner and an outer liner disposed around a combustion space of the turbine engine and extend from an upstream end to a downstream end. The gas turbine engine may also include a plurality of nozzles that extend radially inwards through the outer liner to direct cooling air towards the inner liner. Each nozzle of the plurality of nozzles may extend radially inwards from a first distal end to a second proximal end. Each nozzle of the plurality of nozzles may include multiple air holes arranged in a shower head pattern at the second end.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an illustration of an exemplary disclosed gas turbine engine;

FIG. 2 is a cut-away illustration of an exemplary combustor system of the gas turbine engine of FIG. 1;

FIG. 3 is a cross-sectional view of an embodiment of the outer combustor wall of the gas turbine engine of FIG. 1;

FIG. 4A is a cross-sectional view of another embodiment of the outer combustor wall of the gas turbine engine of FIG. 1;

FIG. 4B is a cross-sectional view of another embodiment of the outer combustor wall of the gas turbine engine of FIG. 1;

FIG. 5A is a cross-sectional view of an embodiment of a nozzle of the gas turbine engine of FIG. 1;

FIG. 5B is another cross-sectional view of the exemplary nozzle of FIG. 5A;

FIG. 6 is a cross-sectional view of another embodiment of a nozzle of the gas turbine engine of FIG. 1; and

FIG. 7 is a cross-sectional view of another embodiment of a nozzle of the gas turbine engine of FIG. 1.

DETAILED DESCRIPTION

FIG. 1 illustrates an exemplary gas turbine engine (GTE) 100 having a compressor system 10, a combustor system 20, a turbine system 70, and an exhaust system 90 arranged lengthwise along an engine axis 98. The compressor system 10 may compress air and deliver the compressed air to an enclosure 72 of the combustor system 20. The compressed air may be mixed with a fuel and directed into a combustor 50 through one or more fuel injectors 30. The fuel-air mixture may ignite and burn in the combustor 50 to produce combustion gases that may be directed to the turbine system 70. The turbine system 70 may extract energy from these combustion gases, and direct the exhaust gases to the atmosphere through the exhaust system 90.

FIG. 2 is a cut-away view of combustor system 20 showing the combustor 50. Combustor 50 includes an outer combustor wall 22 and an inner combustor wall 24 annularly disposed about the engine axis 98. The outer and the inner combustor walls (22, 24) are joined together at an upstream end by a dome assembly to define a combustion space 58 therebetween.

The combustion space 58 is fluidly coupled to turbine system 70 at the downstream end. The plurality of fuel injectors 30, positioned on the dome assembly, direct the fuel-air mixture to the combustion space 58 for combustion. This fuel-air mixture burns in a combustion zone (proximate the upstream end) of the combustion space 58 to produce high pressure combustion gases that flow downstream towards the turbine system 70. The combustion of fuel-air mixture within the combustion space 58 heats the outer and the inner combustor walls (22, 24). For increased reliability and performance of GTE 100, it is desirable to cool these walls. The outer combustor wall 22 includes an inner liner 22 b and an outer liner 22 a, and the inner combustor wall 24 includes an inner liner 24 b and an outer liner 24 a. The inner liners 22 b, 24 b are radially spaced apart from the outer liners 22 a, 24 a to define annular cooling spaces 26, 28 between them. These cooling spaces 26, 28 extend from an upstream end 44 to a downstream end 46 of the combustor 50. The combustion in the combustion space 58 may create oscillations of pressure (pressure waves) within the combustion space 58 that causes radial expansion and contraction (bulging) of the inner liners 22 b, 24 b with respect to the outer liners 22 a, 24 a. The outer liners 22 a, 24 a include a plurality perforations 32, 34 that direct high pressure air from the enclosure 72 to impinge on, and cool, the inner liners 22 b, 24 b. This technology of impingement cooling the combustor liners is referred to in the industry as Augmented Backside Cooled (ABC) technology. It is known that the use of ABC technology decreases the emission of pollutants into the atmosphere. It should be noted that the general configuration of combustor system 20 illustrated in FIG. 2 is exemplary only, and that several variations are possible.

FIG. 3 is a cross-sectional schematic of the outer combustor wall 22 illustrating the impingement cooling of the inner liner 22 b. A high pressure stream of air (“air jets 36”) enters the cooling space 26 through perforations 32 on the outer liner 22 a. These air jets 36 impinge on, and cool, the inner liner 22 b. After impingement, the spent air stream flows towards the downstream end 46 to form the cross-flow air 38 that may be mixed with the combustion gases or discarded. It is known that cross-flow air 38 from the upstream end 44, interacts with, and degrades the ability of the air jets 36 at the downstream end 46 to impinge on, and cool, the inner liner 22 b. For instance, the cross-flow air 38 from a first perforation 32 a may degrade the ability of the air jet 36 from a second perforation 32 b, downstream of the first perforation 32 a, to impinge on the region of the inner liner 22 b under the second perforation 32 b. Similarly, the cross-flow air 38 from the first and second perforations 32 a, 32 b may collectively further degrade the cooling ability of an air jet 36 from a third perforation 32 c, further downstream of the first perforation 32 a, to cool the inner liner 22 b under the third perforation 32 c. In some embodiments, some (or all) of the perforations 32 may include extended ports or nozzles 48 (see FIGS. 4A-7) to reduce the impact of the cross-flow air 38 from an upstream perforation 32 on a downstream air jet 36.

FIG. 4A illustrates a cross-sectional view of the outer combustor wall 22 illustrating nozzles 48 attached to the perforations 32. The nozzles 48 may include air holes 62 that extend from a first end 66, positioned in the enclosure 72 outside the outer liner 22 a, to a second end 68 positioned in the cooling space 26 inside the outer liner 22 a. These air holes 62 may direct the compressed air in the enclosure 72 (air jets 36 of FIG. 3) to impinge on the inner liner 22 b. The nozzles 48 may be a separate part attached to the outer liner 22 a (by any conventional attachment process, such as brazing, etc.) or may be a region of the outer liner 22 a that is bent towards the inner liner 22 b (such as for example, the rim of a perforation that is folded towards the inner liner 22 b). The nozzles 48 may be arranged on the outer liner 22 a such that a radial gap (t) between the second end 68 of a nozzle 48 and the inner liner 22 b decreases from the upstream end 44 to the downstream end 46 (that is, t_(a)>t_(b)>t_(c)>t_(d)>t_(e)). To achieve this decreasing radial gap (t) from the upstream end 44 to the downstream end 46, in some embodiments (as illustrated in FIG. 4A), the length of the nozzles 48 may progressively increase from the upstream end 44 to the downstream end 46. Because the air jets 36 enter the cooling space 26 closer to the inner liner 22 b at the downstream end 46, the effect of the cross-flow air 38 from the upstream air jets 36 on the downstream air jets 36 will be lower. In some embodiments, substantially all the rows of perforations on the outer liner 22 a will include nozzles 48, while in other embodiments, only selected rows of perforations along the length of the outer liner 22 a will include nozzles 48. In some embodiments, the perforations 34 on the inner combustor wall 24 (see FIG. 2) will also include nozzles 48 so that the air jets 36 enter the cooling space 28 closer to the inner liner 24 b at the downstream end 46 than at the upstream end 44.

Although FIG. 4A illustrates the radial gap (t) between the nozzles 48 and the inner liner 22 b as decreasing substantially linearly from the upstream end 44 to the downstream end 46, this is only exemplary. In general, the radial gap (t) may vary in any manner (such as, for example, decrease exponentially from the upstream end to the downstream end). In some embodiments, although the radial gap (t) may generally decrease from the upstream end 44 to the downstream end 46, the radial gaps (t) of selected adjacent nozzles 48 may be substantially the same (such as, for example, t_(a)≈t_(b)>t_(c)>t_(d)≈t_(e)).

In some embodiments, as illustrated in FIG. 4B, only perforations 32 in selected regions of the outer liner 22 a may include nozzles 48 to direct the air jets 36 in these regions closer to the inner liner 22 b. For example, in some embodiments, nozzles 48 may only be included in a few rows of perforations 32 at the downstream end 46 in applications where only the air jets 36 from those few rows are detrimentally affected by the cross-flow air 38 from the upstream end 44. In some embodiments (as illustrated in FIG. 4B), the radial gap (t) between these nozzles 48 and the inner liner 22 b may be substantially the same (that is, t_(a)≈t_(b)). Although FIGS. 4A and 4B illustrate embodiments, where nozzles 48 are used to decrease the radial gap (t) from the upstream end 44 to the downstream end 46, it is contemplated that in some embodiments, the radial gap (t) may instead be decreased by decreasing the distance between the inner liner 22 b and the outer liner 22 a (that is, the thickness of the cooling space 26) from the upstream end 44 to the downstream end 46.

In some applications, the pressure pulses generated in the combustion space 58 during combustion may cause portions of the inner liner 22 b to bulge outwards toward the nozzles 48 in corresponding portions of the outer liner 22 a. Contact between the inner liner 22 b and a nozzle 48 may restrict, or even block, air flow (air jets 36) through the nozzle 48, and result in uneven cooling of the liner. Some embodiments of the nozzles 48 of the current disclosure may be configured to allow the air flow to continue even when they are in contact with the inner liner 22 b.

FIGS. 5A and 5B are cross-sectional illustrations of an exemplary embodiment of a nozzle 48A of the current disclosure. FIG. 5A illustrates a cross-sectional view along a plane parallel to a longitudinal axis 88 of nozzle 48A, and FIG. 5B illustrates a cross-sectional view along a plane transverse to the longitudinal axis 88. In the discussion that follows, reference will be made to both FIGS. 5A and 5B. One or more air holes 62 may direct compressed air out of nozzle 48A at second end 68. In some embodiments, as illustrated in FIG. 5A, the one or more air holes 62 may form a shower head pattern of air holes 62 at the second end 68. In some embodiments, all the air holes 62 may extend from the first end 66 to the second end 68, while in other embodiments (as illustrated in FIG. 5A), a single air hole 62 that extends from the first end 66 may be divided into multiple air holes 62 to form a shower head pattern at the second end 68. The single air hole may be divided into multiple air holes anywhere along the length of nozzle 48A. In some embodiments, as illustrated in FIG. 5A, the single air hole may be divided into multiple air holes proximate the second end 68. In some embodiments, multiple (for example, 2, 3, 4, 5, 6 etc.) air holes 62 may be positioned symmetrically around the longitudinal axis 88 at the second end 68. If the inner liner 22 b, 24 b bulges during operation, the bulging liner may contact a central portion (proximate longitudinal axis 88) of the second end 68 of a nozzle 48. And, since the air holes 62 are distributed around the central portion, some or all of the air holes 62 may remain unblocked by the bulging inner liner 22 b, 24 b. Even if flow through some of the air holes 62 is restricted (or even blocked) by the contacting inner liner 22 b, 24 b, the flow though the remaining air holes 62 may provide sufficient cooling for the inner liner 22 b, 24 b. Thus, a shower head pattern of air holes 62 in nozzle 48A may allow air flow to continue through at least some of the air holes 62 when there is contact between the nozzle 48A and the inner liner 22 b, 24 b.

In some embodiments, nozzle 48A may include one or more projections 74 that project outwards from the second end 68. In some embodiments, at least one of these projections 74 may be located between the outlets of the multiple air holes 62 at the second end 68. Other projections (if any) may be located anywhere on, or proximate, the second end 68. For instance, in some embodiments, the projections 74 may be substantially evenly distributed on the second end 68 of the nozzle 48A. These projections 74 may contact a bulging inner liner 22 b, 24 b and act as a standoff to allow air flow through the air holes 62 of the nozzle 48A. The projections 74 may have any shape and size. For instance, in some embodiments, arc-shaped projections may extend towards the inner liner 22 b, 24 b from the periphery of nozzle 48A. In some embodiments, as illustrated in FIG. 5A, the projections 74 may have a rounded edge to reduce bearing stresses on the inner liner 22 b, 24 b during contact. Although described as a projection, it is contemplated that other features (such as grooves, cut-outs, etc.) that allow air from the air holes 62 to exit out of the nozzle 48A when there is contact between the nozzle 48A and the inner liner 22 b, 24 b, may be provided. In place of, or in addition to, discrete projections 74 on the second end 68, in some embodiments, the shape of a nozzle 48A may include a projecting region on the second end 68. For example, as illustrated in FIG. 6, the second end 68 of an exemplary nozzle 48B may have a curved shape with a projecting central region. In these embodiments, the projecting central region may act as the projection 74 that contacts a bulging inner liner 22 b, 24 b. In addition to this projecting central region, in some embodiments, additional projections 74 may also be provided on second end 68 of nozzle 48B.

In some embodiments (as illustrated in nozzles 48A and 48B of FIGS. 5A-6), central axes (64 a, 64 b, 64 c, 64 d, etc.) of the multiple air holes 62 may be substantially parallel to the longitudinal axis 88. However, in other embodiments, the central axis of an air hole 62 may be inclined with respect to the longitudinal axis 88 of the nozzle. FIG. 7 illustrates an embodiment of a nozzle 48C in with the central axes 64 a, 64 b of the air holes 62 a and 62 b make angles θ_(a) and θ_(b), respectively, with respect to the longitudinal axis 88. The angles θ_(a) and θ_(b) may have the same or different magnitudes. In these embodiments, a bulging inner liner 22 b, 24 b may contact the central portion of the second end 68 and allow air flow through the inclined air holes 62 even in the absence of a projection at the second end 68. In some embodiments, one or more projections 74 may also be provided at the second end 68 of nozzle 48C to act as a stand-off. The inclined air holes 62 may also allow the air flowing through them to diverge and impinge on a larger area of the inner liner 22 b, 24 b.

Any type of nozzle (such as, for example nozzles 48, 48A, 48B, 48C, etc.) may be used in an application. In some applications, a nozzle having one air hole 62 (such as nozzle 48 of FIGS. 4A and 4B) may be used in areas where the possibility of contact with the inner liner 22 b, 24 b is minimal, and a nozzle having multiple inclined air holes 62 (such as nozzle 48C of FIG. 7) may be used where the possibility of contact with the inner liner 22 b, 24 b exists. It should be noted that, although contact between a nozzle 48 and the inner liner 22 b, 24 b is described as being a result of a pressure wave in the combustor 50 that causes a portion of the inner liner 22 b, 24 b to bulge and contact one or more nozzles 48 on the outer liner 22 a, 24 b, this is only exemplary. In some applications, vibration of the combustor 50 may cause contact between the inner liner 22 b, 24 b and the nozzles 48. Contact between a nozzle 48 and the inner liner 22 b, 24 b can occur for various other reasons, and the disclosed system can be used to provide continuous air flow through the nozzles 48 during contact that occurs for any reason.

INDUSTRIAL APPLICABILITY

The disclosed systems and methods of impingement cooling a cylinder liner may be applicable to any turbine engine to reliably and effectively cool the cylinder liner. The disclosed system of impingement cooling is configured to prevent the impingement air flow from being blocked as a result of dimensional changes of the combustor liner during operation of the turbine engine. The operation of a gas turbine engine using a disclosed system of impingement cooling will now be explained.

With reference to FIGS. 1 and 2, during operation of GTE 100, air may be drawn into compressor section 10 and compressed. This compressed air may then be directed to enclosure 72 around the combustor 50. The combustor may enclose a combustion space 58 bounded by a double-walled liner (including inner liners 22 b, 24 b and outer liners 22 a, 24 a). A portion of the compressed air may be mixed with fuel and combusted in the combustion space 58. The combustion heats the inner liners 22 b, 24 b of the combustor 50. A portion of the compressed air in the enclosure 72 is directed though the perforations 32, 34 on the outer liner 22 a, 24 a to impinge on, and cool, the hot inner liner 22 b, 24 b (FIGS. 3). To reduce the impact of cross-flow air 38, from upstream perforations, on the cooling effectiveness of downstream perforations, nozzles 48 are provided on some or all the perforations 32, 34. These nozzles 48 deliver the impingement air jets closer to the inner liner 22 b, 24 b at the downstream end 46 of the combustor 50 and thereby reduce the effect of the cross-flow air on the cooling effectiveness of the downstream air jets. To reduce the possibility of the air jets being blocked by dimensional variations of the liner walls during operation of the turbine engine (such as bulging of the inner liner 22 b, 24 b), the air jets may be provided in a shower head pattern at the tip of the nozzles 48. A shower head pattern of air jets may allow some of the air jets to continue to impinge on, and cool, the inner liner 22 b, 24 b even when a bulging inner liner contacts and blocks some of the air jets.

It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed impingement cooling system and method. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed cooling system. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents. 

What is claimed is:
 1. A gas turbine engine, comprising: an impingement cooled double-walled liner, including an inner liner and an outer liner, disposed around a combustion space of the turbine engine and extending from an upstream end to a downstream end; and a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner, each nozzle of the plurality of nozzles extending radially inwards from a first distal end to a second proximal end, the plurality of nozzles being arranged such that a radial gap between the second end of a nozzle and the outer liner decreases from the upstream end to the downstream end, wherein, at least one nozzle of the plurality of nozzles includes multiple air holes at the second end.
 2. The gas turbine engine of claim 1, wherein the at least one nozzle further includes a longitudinal axis extending from the first end to the second end and each air hole of the multiple air holes includes a central axis, the multiple air holes being symmetrically arranged about the longitudinal axis.
 3. The gas turbine engine of claim 2, wherein the central axis of each air hole is substantially parallel to the longitudinal axis.
 4. The gas turbine engine of claim 2, wherein the central axis of each air hole is inclined with respect to the longitudinal axis such that the cooling air exiting the at least one nozzle diverges.
 5. The gas turbine engine of claim 1, wherein the multiple air holes in the at least nozzle is arranged in a shower head pattern at the second end.
 6. The gas turbine engine of claim 1, wherein the second end of the at least one nozzle includes a projection that extends towards the inner liner.
 7. The gas turbine engine of claim 6, wherein the projection is centrally positioned on the second end and each air hole of the multiple air holes is symmetrically positioned about the projection.
 8. The gas turbine engine of claim 1, wherein the second end of the at least one nozzle is curved such that a central portion of the second end forms a proximal-most portion of the nozzle.
 9. The gas turbine engine of claim 1, wherein the radial gap decreases substantially linearly from the upstream end to the downstream end.
 10. A method of impingement cooling a double-walled combustor liner of a gas turbine engine, the double-walled liner extending from an upstream end to a downstream end and including an inner liner and an outer liner positioned radially outwards the inner liner, comprising: combusting a fuel in a combustor of the gas turbine engine; and directing cooling air through a plurality of nozzles extending radially inwards through the outer liner to impinge upon and cool the inner liner, such that the cooling air exits the plurality of nozzles closer to the inner liner at the downstream end than at the upstream end, wherein the cooling air directed through at least one nozzle of the plurality of nozzles exit the at least nozzle through multiple air flow paths symmetrically arranged about a longitudinal axis of the at least one nozzle.
 11. The method of claim 10, wherein directing the cooling air includes directing the cooling air through the multiple air flow paths of the at least one nozzle such that the cooling air diverges.
 12. The method of claim 10, wherein directing the cooling air includes directing the cooling air though the multiple air flow paths of the at least nozzle such that the cooling air through each of the multiple air flow paths flow substantially parallel to one another.
 13. A gas turbine engine, comprising: an impingement cooled double-walled liner, including an inner liner and an outer liner, disposed around a combustion space of the turbine engine and extending from an upstream end to a downstream end; and a plurality of nozzles extending radially inwards through the outer liner to direct cooling air towards the inner liner, each nozzle of the plurality of nozzles extending radially inwards from a first distal end to a second proximal end, wherein each nozzle of the plurality of nozzles include multiple air holes arranged in a shower head pattern at the second end.
 14. The gas turbine engine of claim 13, wherein the plurality of nozzles are arranged such that a radial gap of the second end of a nozzle to the inner liner decreases as a function of distance from the upstream end to the downstream end.
 15. The gas turbine engine of claim 13, wherein the multiple air holes are symmetrically positioned about a longitudinal axis of each nozzle.
 16. The gas turbine engine of claim 15, wherein each air hole of the multiple air holes are inclined with respect to the longitudinal axis such that the cooling air exiting each nozzle diverges.
 17. The gas turbine engine of claim 16, wherein an inclination of each air hole of the multiple air holes with respect to the longitudinal axis is substantially the same.
 18. The gas turbine engine of claim 13, wherein the second end of each nozzle includes a projection that extends towards the inner liner.
 19. The gas turbine engine of claim 18, wherein the multiple air holes are symmetrically arranged about the projection.
 20. The gas turbine engine of claim 13, wherein the second end of each nozzle is curved such that a central portion of the second end forms a proximal-most portion of the nozzle. 